Gas turbine engine airflow member having spherical end

ABSTRACT

A gas turbine engine airflow member is disclosed having spherical interface shapes at a tip end and a hub end. In one embodiment, the airflow member is a fan blade. The spherical interface shapes can be convex or concave. In one form, the tip end spherical shape is axially shifted aft relative to a hub end. The tip end can include relatively little forward portion in an expanding flow path with predominant portion of the tip end in a contracting flow path. Such a configuration can occur with a midpoint and trailing edge located in the contracting flow path. A center of gravity of a swept airflow member having spherical end shape can be located on a pivot axis of the airflow member. Divot depressions can be provided to permit concave flow path shapes.

CROSS REFERENCE TO RELATED APPLICATIONS

This application claims priority to and the benefit of U.S. ProvisionalPatent Application No. 61/775,635, filed 10 Mar. 2013, the disclosure ofwhich is now expressly incorporated herein by reference.

TECHNICAL FIELD

The present disclosure generally relates to gas turbine engine airflowmembers having tip clearance features. More particularly, but notexclusively, the present disclosure relates to tip clearance usingspherical shapes.

BACKGROUND

Providing consistent and small tip clearance in gas turbine engineairflow members remains an area of interest. Some existing systems havevarious shortcomings relative to certain applications. Accordingly,there remains a need for further contributions in this area oftechnology.

SUMMARY

One embodiment of the present disclosure is a unique gas turbine engineairflow member. Other embodiments include apparatuses, systems, devices,hardware, methods, and combinations for maintaining small tip clearancesover a range of operating conditions of a gas turbine engine airflowmember. Further embodiments, forms, features, aspects, benefits, andadvantages of the present application shall become apparent from thedescription and figures provided herewith.

BRIEF DESCRIPTION OF THE FIGURES

FIG. 1 depicts one embodiment of a gas turbine engine;

FIG. 2 depicts one embodiment of an airflow member;

FIG. 3 depicts another embodiment of an airflow member;

FIGS. 4A and 4B depict embodiments of an airflow member;

FIG. 5 depicts an embodiment of a divot depression;

FIGS. 6A and 6B depict embodiments of divot depressions;

FIG. 7 depicts various embodiment of hub interface shapes; and

FIG. 8 depicts various embodiments of tip interface shapes.

DETAILED DESCRIPTION OF THE ILLUSTRATIVE EMBODIMENTS

For the purposes of promoting an understanding of the principles of thedisclosure, reference will now be made to the embodiments illustrated inthe drawings and specific language will be used to describe the same. Itwill nevertheless be understood that no limitation of the scope of thedisclosure is thereby intended. Any alterations and furthermodifications in the described embodiments, and any further applicationsof the principles of the disclosure as described herein are contemplatedas would normally occur to one skilled in the art to which thedisclosure relates.

With reference to FIG. 1, one embodiment of a gas turbine engine 50 isdepicted which includes a fan 52, a compressor 54, a combustor 56, and aturbine 58. Air is received into and compressed by the compressor 54prior to being delivered to the combustor 56 where it is mixed with fueland burned. A flow of air and products of combustion is then deliveredto the turbine 58 which expands the flow stream and produces work thatis used to drive the compressor 54 as well as to drive the fan 52. Thefan 52 is used to develop thrust by accelerating air through a bypasspassage 60 which is exhausted out of the rear of the engine 50.

The gas turbine engine can be used to provide power to an aircraft andcan take any variety of forms. As used herein, the term “aircraft”includes, but is not limited to, helicopters, airplanes, unmanned spacevehicles, fixed wing vehicles, variable wing vehicles, rotary wingvehicles, unmanned combat aerial vehicles, tailless aircraft, hovercrafts, and other airborne and/or extraterrestrial (spacecraft)vehicles. Further, the present disclosures are contemplated forutilization in other applications that may not be coupled with anaircraft such as, for example, industrial applications, powergeneration, pumping sets, naval propulsion, weapon systems, securitysystems, perimeter defense/security systems, and the like known to oneof ordinary skill in the art.

Though the engine 50 is depicted as a single spool engine, otherembodiments can include additional spools. The embodiment of the engine50 depicted in FIG. 1 is in the form of a turbofan engine, but it willbe appreciated that some embodiments of the gas turbine engine can takeon other forms such as, but not limited to, open rotor, turbojet,turboshaft, and turboprop. In some forms, the gas turbine engine 50 canbe a variable cycle and/or adaptive cycle engine.

Turning now to FIG. 2, an airfoil member 62 that can be used in theturbomachinery components of the gas turbine engine 50 is depicted. Theairfoil member 62 is an airfoil shaped elongate component that extendsacross a flow path of the turbomachinery component and which can be usedto operate upon a fluid traversing the flow path, such as by changing adirection and/or pressure of the fluid travelling through the flow path.

The embodiment of the airfoil member 62 depicted in FIG. 2 is in theform of a rotatable blade capable of being rotated around the centerline64, but in other embodiments it will be appreciated that the airfoilmember 62 may take the form of the stator vane. The airfoil member 62 isdisposed in a flowpath 66 formed between an inner wall 68 and an outerwall 70 and is capable of being rotated to a pitch angle about the axis72. The flowpath 66 can be an annular flow path in many embodiments. Theairfoil member includes a tip end 74 disposed adjacent the outer wall70, and a hub end 76 disposed adjacent the inner wall 68.

The inner wall 68, outer wall 70, tip end 74, and hub end 76 areillustrated for purposes of convenience as straight lines, but as willbe discussed further below one or more of the walls 68, 70 and ends 74,76 will have a spherical shape to provide for a substantially constantclearance over a wide range of pivoting action of the airfoil member 62.To set forth just one non-limiting example, the outer wall 70 and tipend 74 can have complementary spherical shapes, while the inner wall 68and hub end 76 do not. It is also possible for both outer and innerwalls 70, 68 to have spherical shapes. As used herein, the term“spherical” will be understood to include those shapes that aresubstantially spherical to provide for a substantially constantclearance over the wide range of pivoting the action of the airfoilmember 62. To set forth just a few non-limiting examples, some variationin the shape of the interface will be recognized for manufacturingtolerances and service wear.

The spherical shape at the tip and hub ends of the airfoil member 62 canbe concave or convex on either or both tip and hub. For example, in theembodiment depicted in FIG. 2, the tip end of the airfoil member 62 canhave a convex shape such that the flow path extends up and away from thehub end 76 at a leading edge 78 of the airfoil member before reaching anapex and descending back towards the hub and 76 at a trailing edge 80 ofthe airfoil member 62. Various other possibilities will be appreciatedfrom this singular example. For example, the tip end 74 can have aconcave shape, the hub end 76 can have a convex shape, and the hub end76 in other embodiments can have a concave shape. Thus, any variation ofconcave and convex shapes can be used on either or both of the tip end74 and hub end 76. Furthermore, the ends 74, 76 can both be either aconcave or convex shape, while in other embodiments one of the ends 74,76 can be concave while the other of the ends 74, 76 can be convex.

The relative placement of the concave and convex shape will determinewhether the shape is extending away from or toward the centerline 64.For example, if the outer wall 70 is concave, part of that surface mayextend away from the centerline while another part is extending towardthe centerline. Such a situation is present when the outer wall 70 is atthe same radial distance from the centerline 64 on both sides of theconcave spherical shape.

Separate and apart from the shape of the walls 68 and 70 and whether ornot the walls are extending toward or away from the centerline 64, thearea of the flow path annulus can also be increasing or decreasing duein large part to the interdependent nature of flow path area with outerradius and inner radius. In various embodiments, the cross sectionalarea may be increasing or decreasing as a function of axial locationdepending on the relative nature of the spherical shapes at one or bothof the walls 68 and 70. No limitation on the nature of the crosssectional area is intended unless expressly stated to the contrary.

Though the airfoil member 62 is shown for convenience as aquadrilateral, it will be appreciated that the airfoil member 62 canhave any variety of characteristics associated with any number ofairfoil members used in gas turbine engines. For example, the airfoilmember 62 can include sweep, lean, twist, etc. Furthermore, the airfoilmember 62 can have any variety of orientations of its characteristics(sweep, lean, twist, etc) relative to any variety of features (e.g.pivot axis, leading and trailing edges, etc) of the airfoil member 62.For example, the tip end 74 and the hub end 76 can be located relativeto the axis 72 such that a greater proportion of either of the tip and74 or the hub end 76 is disposed either forward or aft of the axis 72.Other embodiments will be described further below.

The spherical shape can occur in any range between the leading edge 78and trailing edge 80 including the entire range between leading andtrailing edge. In some forms, a spherical shape can be located betweenthe leading edge and a point between the leading edge and trailing edge(e.g. a pivot point between the leading edge and trailing edge), andanother spherical shape with a different radius can be located betweenthe midpoint and the trailing edge. In fact, the different sphericalshapes can be a mixture of concave and convex spherical shapes. To setforth just one non-limiting example, the forward spherical shape can beconcave and the aft spherical shape can be convex. FIGS. 7 and 8 setforth a few non-limiting examples of having spherical shapes forward andaft of the pivot axis, whether the shapes are a mixture of concave,convex, or wholly concave or convex. The shapes can have the same ordifferent radii. The arrow indicates rotation about the axis in eitherdirection. The various figures also illustrate a combination ofspherical shapes and linear shapes that also produce very little gapwhen the member is pivoted about the axis. The straight line, or linear,segments can be either forward or aft of the pivot axis, or both forwardand aft. It will be appreciated that any particular hub side geometrydepicted in any of the nine separate variations of FIG. 7 can be pairedwith any particular tip side geometry depicted in any of the nineseparate variations in FIG. 8.

Turning now to FIG. 3, one embodiment of the airfoil member is shown asa fan blade 62 capable of pivoting about axis 72 and rotatable about thecenterline 64. The flowpath 66 is bounded by a hub that generallyextends away from the centerline 64 at an upstream end until reaching anapex before descending towards the centerline 64. The fan blade 62 isdepicted as being located near an apex of the hub, but in other formsthe fan blade 62 can be located further forward on the hub or furtheraft.

Turning now to FIGS. 4A and 4B, two separate embodiments of the fanblades 62 are depicted, both of which include a spherical tip end 74.Each of the fan blades 62 illustrated in FIG. 4A and FIG. 4B includecompound sweep distributions with the primary difference being that theembodiment in FIG. 4A includes a rearward sweep near the hub end 76 anda forward sweep at the tip end 74, while the embodiment and FIG. 4Binitially includes a rearward sweep at the hub followed by successiveforward sweep, rearward sweep, and finally forward sweep at the tip.Another difference is that the axis 72 of the embodiment depicted inFIG. 4B passes through a center of gravity 82 of the fan blade 62. Whileboth embodiments depict compound sweep distributions, other forms of thefan blade 62 good consist of backward sweep only.

Turning now to some similarities between FIG. 4A and FIG. 4B, it will benoted that the tip in 74 is shifted axially aft relative to the hub and76. For example, the leading edge of the tip end 74 is shifted backaxially relative to the leading edge of the hub in the 76. Thearrangement depicted in FIG. 4A and FIG. 4B also permits the sphericalshape at the tip end 74 to have mostly contraction which permits higherfan pressure ratios for a given diameter relative to a flowpath withlesser contraction. A descending flow path can also minimize enginelength, and thus engine weight, because the S-duct downstream of the fanblade 62 can be shorter and the core vane, sometimes referred to as theESS, can be on a descending flow path. Having the ESS on the descendingflow path can reduce the extreme radius change in the S-duct given thatthe starting radius is reduced.

The spherical shape at the tip end 74 includes an apex at the axis 72.Given that the pivot axis 72 is located towards the leading edge 78forward of a midpoint between the leading edge 78 and the trailing edge80, the portion of the spherical shape located between the leading edge78 and the pivot axis 72 is smaller relative to the portion of thespherical shape located between the pivot axis 72 and the trailing edge80. The midpoint 84 can represent half the arc distance of the sphericalshape at the tip end 74 between the leading edge 78 and trailing edge80.

In one embodiment of the fan blade 62 illustrated in FIGS. 4A and 4B,several aerodynamic considerations can be taken into account improveoperation. For example, the fan hub pressure ratio can be limited toavoid separating the flow at the trailing edge 80 of the fan blade 62.To accommodate this, all of the turning of the flow in the hub regionnear the leading edge 78 can be accomplished and negative camber can beincluded in the airfoil to keep airflow attached at the trailing edge.

FIG. 5 depicts an embodiment of the inner wall 68 in the form of aturbofan spinner having a divot depression 86 formed downstream of a tip88 of the spinner. The divot depression 86 can include a spherical shapeas will be appreciated given some of the discussion above. The divotdepression 86 is formed upstream of an apex 90 associated with theturbofan flowpath, but it will be appreciated that the divot depression86 can be formed in any variety of locations including at the apex 90and downstream of it. The divot depression 86 includes a spherical shapesize to receive a complementary circle shaped hub end 76 of a fan blade62 (not shown). Also depicted in FIG. 5 is a pivot point 92 throughwhich the axis 72 extends. The pivot 92 is shown located in a relativelyforward portion of the divot depression 86, but in other embodiments thepivot 92 can be placed in other locations of the divot depression 86. Inaddition, the axis 72 is illustrated at a right angle to the centerline64, but in other embodiments the axis 72 can be oriented at other anglesrelative to the centerline 64.

Turning now to FIGS. 6A and 6B, two separate embodiments are shown of agas turbine engine having a plurality of divots 86 disposedcircumferentially about an annular flow path. The embodiment depicted inFIG. 6A illustrates divot depressions 86 separated by a distance thatcan be referred to as a pitch. The distance, or pitch, between divotdepressions 86 can be determined based upon solidity of airfoil members62, among other things. Though the divot depressions 86 are not limitedto any particular form of the airflow member 62, in one non-limitingembodiment the divot depressions 86 are used with a fan blade 62 at thehub location that suggested above in FIG. 5.

FIG. 6B illustrates divot depressions 86 that are spaced close enoughtogether such that the divot depressions 86 overlap one another. Thedashed lines depicted in FIG. 6B illustrate what would be nominalcontour lines of respective divot depressions 86 were it not for theoverlap with a neighboring divot depression 86. Overlapping divotdepressions 86 may occur when solidity is greater than 1.0. Two separateangles are also illustrated in FIG. 6B. An angle α₁ illustrates an anglemeasured from the pivot 92 between a reference axis 94 and a locationthat a leading edge 78 of the airflow member 62 would intersect aforward divot overlap between adjacent divot depressions 86. Likewise,angle α₂ illustrates an angle measured from the pivot 92 between areference axis 94 and a location that a trailing edge 80 of the airflowmember 62 would intersect an aft divot overlap between adjacent divotdepressions 86. The reference axis 94 can be any axis useful to measuredivot overlaps, and, in one form, is the centerline 64. The angles α₁and α₂ can be calculated based on solidity, blade count, and pivot axislocation to ensure that certain performance critical points have tightclearances between the airfoil member 62 and a surface of the divotdepression 86. Operation of the airflow member 62 at a location in whichat least portion of it extends past the divot overlap angles is stillacceptable at other performance points where inefficiency may be lessimportant. For example, reverse pitch operation for generation ofreverse thrust may be an example of operation of the airflow member 62in which it extends past the divot overlap angle. The angles α₁ and α₂can be the same in some embodiment, but will be different in others asshown in the illustrated embodiment. In one particular form, the anglesα₁ and α₂ are different when the pivot point 92, and corresponding axis72, are placed forward of a mid-chord location of the airflow member 62.

Any of the embodiments disclosed herein of the particular arrangement ofthe tip can be used with any of the embodiments disclosed herein of theparticular arrangement of the hub. For example, axially displacing thetip end 74 relative to the hub in 76 where the tip in 74 includes mostlycontraction, and combining that arrangement with the hub end 76 in whichthe pivot axis is placed in front of the mid chord location at the hubend 76. This combination can include concave or convex vertical shape atthe hub end 76. Any variety of other combinations are also contemplated.

While the invention has been illustrated and described in detail in thedrawings and foregoing description, the same is to be considered asillustrative and not restrictive in character, it being understood thatonly the preferred embodiments have been shown and described and thatall changes and modifications that come within the spirit of thedisclosures are desired to be protected. It should be understood thatwhile the use of words such as preferable, preferably, preferred or morepreferred utilized in the description above indicate that the feature sodescribed may be more desirable, it nonetheless may not be necessary andembodiments lacking the same may be contemplated as within the scope ofthe disclosure, the scope being defined by the claims that follow. Inreading the claims, it is intended that when words such as “a,” “an,”“at least one,” or “at least one portion” are used there is no intentionto limit the claim to only one item unless specifically stated to thecontrary in the claim. When the language “at least a portion” and/or “aportion” is used the item can include a portion and/or the entire itemunless specifically stated to the contrary.

Unless specified or limited otherwise, the terms “mounted,” “connected,”“supported,” and “coupled” and variations thereof are used broadly andencompass both direct and indirect mountings, connections, supports, andcouplings. Further, “connected” and “coupled” are not restricted tophysical or mechanical connections or couplings.

What is claimed is:
 1. A gas turbine engine comprising an airflow memberdisplaced from an engine centerline and disposed in a gas path between aradially outer flow path boundary and a radially inner flow pathboundary, the airflow member having a tip end spaced apart from theradially outer flow path boundary that together form a tip interface,wherein the tip interface includes a spherical interface shape having aspherical radius that extends over an axial range and is characterizedby a forward axial end, an axial midpoint, and an aft axial end, theaxial midpoint and aft axial end located in a region in which thespherical interface shape is descending toward the engine centerline. 2.The gas turbine engine of claim 1, wherein the airflow member includes ahub end spaced apart from the radially inner flow path boundary and thetip end is shifted axially aft relative to the hub end.
 3. The gasturbine engine of claim 2, wherein the forward axial end is located in aregion in which the flow path is expanding away from the enginecenterline.
 4. The gas turbine engine of claim 3, wherein the airflowmember includes a sweep and a center of gravity of the airflow member islocated on the adjustment axis.
 5. The gas turbine engine of claim 3,wherein the airflow member is a variable rotor blade having anadjustment axis, the variable rotor blade is pivotable about theadjustment axis, and a trailing edge at a hub end of the variable rotorblade includes a negative camber.
 6. The gas turbine engine of claim 5,wherein the adjustment axis is in front of a mid chord location at thetip end.
 7. The gas turbine engine of claim 4, wherein the airflowmember includes a forward sweep.
 8. The gas turbine engine of claim 1,wherein the airflow member further includes a hub end near the radiallyinner flow path boundary that together form a hub interface, the airflowmember is pivotable about a pivot axis such that the tip end and the hubend are rotatable relative to the radially outer flow path boundary andradially inner flow path boundary, and the hub interface includes a hubspherical interface shape having a spherical radius.
 9. The gas turbineengine of claim 1, the radially inner flow path boundary at the aft endof the airflow member descends towards an engine centerline.
 10. An gasturbine engine comprising a variable position airfoil member thatincludes a hub end having a leading edge and a trailing edge, thevariable position airfoil member disposed within a flow path annulusadjacent a hub flow path surface having a divot that includes adownstream end in which a radial rate of growth of the downstream end isincreasing, the variable position airfoil member pivotable around anadjustment axis and capable of being moved between a first position anda second position where the first position provides a flow area greaterthan a flow area in the second position, the hub end offset from thedivot to permit pivoting movement of the variable position airfoilmember without interference from the flow path surface between the firstposition and the second position, wherein the offset between the flowpath surface of the divot and the hub end of the variable positionairfoil member is constant as the variable vane moves between the firstposition and the second position.
 11. The gas turbine engine of claim10, which further includes a plurality of variable position airflowmembers associated with a plurality of divots distributed around thecircumferentially extending flow path surface, and wherein a pluralityof adjustment axes associated with the plurality of variable airflowmembers intersect at a common origin, and wherein the hub end has agreater proportion disposed forward of the adjustment axis relative to aproportion of the tip end disposed forward of the adjustment axis. 12.The gas turbine engine of claim 11, wherein the variable positionairfoil member is pivotable between an upper limit position beyond thefirst position and the offset between the first and second position isdifferent than the offset between the upper limit position and the firstposition.
 13. The gas turbine engine of claim 11, wherein a solidity ofthe variable position airfoil members is less than one and the pluralityof divots are separated by a hub surface pitch.
 14. The gas turbineengine of claim 10, wherein the hub end is spherical, the variableposition airflow member includes a tip end that is spherical, and theaxial extent of spherical shape at the tip end forward of a pivot axisof the variable position airflow member is shifted backward relative tothe axial extent of the spherical shape at the hub end forward of thepivot axis.
 15. The gas turbine engine of claim 14, wherein one of thetip offset and hub offset includes a first range of pivot angles suchthat clearance is substantially constant and a second range of pivotangles such that clearance in the second range is variable.
 16. The gasturbine engine of claim 15, wherein a forward portion of the variableposition airflow member includes a first forward range of pivot anglesover which the offset is constant and an aft portion of the airflowmember includes a first aft range of pivot angles over which the offsetis constant and the first forward range is different than the first aftrange.
 17. The gas turbine engine of claim 10, wherein the adjustmentaxis is biased to one side of the divot.
 18. A method comprisingrotating a turbomachinery blade about a centerline of a gas turbineengine, the turbomachinery blade having a spherical shaped end, pivotingthe turbomachinery blade about a pivot axis to swing from a firstposition to a second position, the spherical shaped end residing in aconcave divot depression formed in a flow path surface, the concavedivot depression having a spherical shape complementary to the sphericalshaped end of the turbomachinery blade, as a result of the pivoting,sweeping the spherical shaped end of the turbomachinery blade across thecomplementary spherical shape of the concave divot depression, andmaintaining a substantially constant clearance between the sphericalshaped end of the turbomachinery blade and the spherical shape of theconcave divot depression.
 19. The method of claim 18, wherein thepivoting results in extending a trailing edge of the turbomachineryblade past the divot depression after the turbomachinery blade reaches afirst limit angle.
 20. The method of claim 19, wherein the pivotingresults in extending a leading edge of the turbomachinery blade past thedivot depression after the turbomachinery blade reaches a second limitangle.